The present invention relates to turbines, and more particularly to gas turbines with a liquid cooling system encapsulated into the turbine rotor.
The present invention further relates to a liquid cooling system which includes a tank containing a cooling liquid and a coolant pump coupled to the tank for forced recirculation of the coolant liquid along the channels formed in the rotor shaft and through the rotor disks and blades of the turbine, and wherein the coolant pump is actuated as a result of relative motion of the stator with regard to the rotor of the turbine.
The present invention also relates to a turbine liquid cooling system which includes a heat exchanger positioned in a compressor section of the turbine, either in a compressor drum, or at the end of the compressor section, or the combination thereof, and wherein, being positioned at the end of the compressor section, the heat exchanger may use the blade thereof or the blade of the compressor for cooling purposes.
As known to those skilled in the art, a gas turbine is a heat engine that converts a portion of the fuel energy into work by using gas as the working medium which commonly delivers its mechanical output through a rotating shaft. In the typical gas turbine, the sequence of thermodynamic processes consists basically of compression, addition of heat in a combustor, and expansion through a turbine. This basic operation of the gas turbine may be modified through the addition of heat exchangers and multiple components for reasons of efficiency, power output, and operating characteristics. In order to achieve an overall performance of the turbine, each process is carried out in the engine by a specialized component.
As shown in FIG. 1A, a typical turbine 1 includes: a rotor 2 having a rotating shaft 3 with disks 4 and blades 5, attached thereto, a stator 6, as well as a combustor 7, and a compressor 8.
In operation, air for the combustion chamber 7 is forced into the turbine by the compressor 8. In the combustor 7, fuel is mixed with the compressed air and is burned. Heat energy, thus released, is converted by the turbine into rotary energy.
High emission temperature of the combustion product generally necessitates excess air to cool the combustion product to the allowable turbine inlet design temperature. To improve efficiency, heat exchangers may be added to the gas turbine exhaust to recover heat energy and return heat to the working medium after compression and prior to heat combustion.
Two basic types of compressors are used in gas turbines, namely, axial and centrifugal. In a few special cases, a combination type known as a mixed wheel, which is partially centrifugal and partially axially, is used.
Combustors, sometimes referred to as combustion chambers, for gas turbines take a wide variety of shapes and forms. All contain nozzles to meter the fuel to the gas stream and to atomize or break up the fuel stream for efficient combustion. In addition to being designed to burn the fuel efficiency, the combustors also uniformly mix excess air with the products of combustion to maintain a turbine at a uniform lower temperature. The combustor brings the gas to a controlled uniform temperature with a minimum of impurities and a minimum loss of pressure.
The turbine itself includes the rotor 2 with turbine disks 4 and blades 5 thereon, and a stopper. Two types of gas turbine disks are generally used, namely, radial-in-flow and axial-flow. Small gas turbines usually use a radial flow disk, while for larger volume flows, axial turbine disks are used almost exclusively. Although some of the turbines are of the simple impulse type, most high performance turbines are neither pure impulse nor pure reaction. The high performance turbines are normally designed for varying amounts of reaction and impulse to the optimum performance.
Gas turbines are typically provided with subsystems of control and fuel regulation. The primary function of the subsystem that supplies and controls the fuel is to provide clean fuel, free of vapor, at a rate appropriate to engine operation conditions. These conditions may vary rapidly and over a wide range. As a consequence, fuel controls for gas turbines are, in effect, special purpose computers employing mechanical, hydraulic, or electronic means, with all three in combination being frequently used.
All gas turbines employ some kind of cooling to various extent and use a liquid or gas coolant to reduce the temperature of the metal parts. The cooling system varies from the simplest form where only first stage disk cooling is involved to the more complex systems where the complete turbine (rotor, stator and blades) is cooled. Two basic types of heat exchangers are used in gas turbines, namely, gas-to-gas and gas-to-liquid. An example of the gas-to-gas type is the regenerator which transfers heat from the turbine to the air leaving the compressor. The regenerator must withstand rapid large temperature changes and must have low-pressure drop. Intercoolers, which are used between stages of compression are generally air-to-liquid units. They reduce the work of compression and the final compressor discharge temperature. When used with a regenerator, they increase both the capacity and efficiency of a gas turbine of a given size.
The thermal efficiency of gas turbines is critically dependent on the temperature of burnt gases at the turbine inlet. The higher temperature generally results in a higher efficiency. Stochiometric combustion would provide maximum efficiency, however in the absence of an internal cooling system, turbine blades cannot tolerate gas temperatures that exceed 1300 K. For this temperature, the thermal efficiency of turbine engine is only 52%. Conventional air-cooling techniques of turbine blades allow inlet temperatures of about 1500 K on current operating engines yielding thermal efficiency gains of about 4%. Newer designs, that incorporate advanced air-cooling methods allows inlet temperatures of 1750-1800 K, with a thermal efficiency gain of about 3.5% compared to current operating engines. This temperature is near the limit allowed by air-cooling systems.
Turbine blades may be cooled with air taken from the compressor or by liquid. Cooling systems with air are easier to design but have a relatively low heat transfer capacity and reduce the efficiency of the engine. Some cooling systems with liquid rely on thermal gradients to promote re-circulation from the tip to the root of turbine blades. In these cases, the flow and cooling of liquid are restricted. For optimum results, cooling systems with liquid (shown in FIG. 1B) should use a pump 9 to recirculate the coolant 10 contained in coolant tank 11 and a heat exchanger 12 to cool the liquid. In the past, designers have tried to locate the pump 9 on the engine stator 6 and, therefore were unable to avoid high coolant losses through seals 13xe2x80x2 and 13xe2x80x3 since it has been found to be impossible to reliably seal the stator-rotor interface.
The Carnot cycle provides the theoretical limit for the thermal efficiency of any heat engine. This limit, xcex7Carnot, is given by                               η          carnot                =                  1          -                                    T              L                                      T              H                                                          (        1        )            
where TL is the absolute temperature of the low-temperature reservoir, and TH is the absolute temperature of the high-temperature reservoir.
For gas turbine engines, maximum thermal efficiency,                               η          1                      (            max            )                          =                  1          -                                                    T                1                *                                            T                3                *                                                                        (        2        )            
where T1* and T3* are the total temperature at compressor and turbine inlet, respectively. Therefore, increasing the temperature at the turbine inlet is the most advantageous method for improving the efficiency and power of gas turbine engines. Simultaneously, the specific weight and frontal area of the engine decrease. The improvement of these two performance parameters is especially important for aeroengines.
The design of turbine engines has been continuously perfected for many years. Newer engines have been designed which are more powerful and more efficient as the need has arisen. Nevertheless, during the last twenty years, engine designers have been increasingly held back by the law of diminishing returns. On the average, despite sustained efforts, from 1970 to 1997, T3* has increased by only 10K per year.
The control of blade temperature is an important aspect of turbine cooling. Turbine vanes are fixed and therefore can be easily cooled using known methods (including liquid systems) and are subject to little mechanical stress. In contrast, rotor blades are simultaneously subject to high levels of both mechanical and thermal stress. The mechanical stress is due to gas pressure, high centrifugal force and vibration. Thermal stress is due to heat transfer from burnt gases and is especially severe at the leading and trailing edge of the rotor blade. A blade is designed to work at the mechanical limit of the material. For this reason, in parallel with research on material science, great efforts have been made to improve the cooling of rotor blades. Most performance parameters (such as thermal efficiency and thereby overall engine efficiency, specific fuel consumption, weight, thrust and frontal area of the engine) depend on T3*. This temperature is limited by the heat transfer capacity of the system used for cooling the turbine blades. Therefore, it is an important design consideration to improve this cooling system.
Most current systems use air as a coolant. There are three basic air-cooling systems, namely, convection, film, and transpiration cooling. In convection cooling, cold air is brought from compressor through openings. This air cools the turbine disc, then circulates through holes or openings within the blade toward the blade tip and exits through the leading and trailing edge or through the blade tip. Due to mechanical and manufacturing constraints, blades have a limited internal cooling surface and air is generally not an efficient cooling fluid. When the temperature at turbine inlet becomes very high, the amount of air required for cooling increases to an unacceptable level. Film cooling systems blast cold air on the external surface of turbine blades. A drawback of this system is the difficulty of controlling the flow of cold air around the blade surface. Therefore, designers generally agree that further research on air based cooling systems would yield only marginal T3* gains. The maximum temperature at turbine inlet provided by air-cooling systems is about 1800 K and can be sustained only for a limited time.
Currently, transpiration (effusion) cooling is to a great extent the method of choice. This system attempts to isolate turbine blades from burnt gases using a thin layer of low temperature air. In this case, blades must be coated with a porous material in order to allow the air efflux. Therefore this cooling system faces daunting challenges. The porous material is subject to excessive wear, fatigue and corking and has to be internally cooled by the same airflow. The porous coating has a reduced strength, while the maximum mechanical and thermal stress is reached on the blade surface. The air film disturbs the flow pattern of burnt gases leading to power losses. Due to mechanical vibrations and combustion processes, the flow of burnt gases is turbulent, therefore, the thin layer of relatively cold air is difficult to maintain. The porous layer is further exposed to erosion caused by flow and corrosive attack due to reactive gases present in burnt gases, especially sulfur compounds and oxidants. In time, these phenomena reduce the strength, thickness and continuity of the porous layer. In addition, burnt gases contain a degree of impurities. These impurities are generally in a plastic state due to the high gases temperature and occlude the external pores of the blade. As a result, cooling is greatly reduced.
Liquid cooling systems have been tested from the beginning of the design of turbines, together with air-based systems. As late as the 1950s, liquid cooling systems were regarded as having the greatest potential. In general, engine designers have placed the liquid tank, pump and heat exchanger on the engine stator. Referring to FIG. 1B, the cooling liquid 10 is aspired from the tank 11, then directed through a duct to the first stator/rotor sealing 13xe2x80x2. Through this seal, the liquid enters the engine rotor 2, initially cools the turbine disc 4, passes through radial holes into the blades 5, and returns to the engine stator 6 through a second seal 13xe2x80x3. Special ducts direct the hot liquid that has returned from the rotor 2 to the heat exchanger 12. Cool liquid exits from the heat exchanger 12 and returns to the tank 11 with the cycle being repeated. Distillated water was the liquid used during early experiments. Water was vaporized within turbine blades 5 and therefore the second seal 13xe2x80x3 did not prevent excessive coolant losses. Other significant losses occurred through the first seal 13xe2x80x2. The high volume of the heat exchanger 12 increased the frontal area of the engine, a feature that reduces aircraft performance. For these reasons, cooling systems based on forced recirculation of liquid have fallen by the wayside.
Open loop systems that spray liquid from the stator vanes on the blades have been investigated. The open loop systems require a prohibitive amount of coolant (especially for aero-engines), cool incompletely and are difficult to control. Recently, General Electric designed an open loop cooling system with steam. Unfortunately this system is limited to terrestrial power applications.
For an extended period of time, design engineers were attracted by promoting liquid recirculation using density variations (thermosyphon cooling). This concept has basic design consideration problems. Due to space constraints, the heat exchanger has a small surface and the presence of ribbed legs reduces the disc strength. Nevertheless, this concept does contain the seed of a successful solution, since the engine rotor contains the entire cooling system thereby stator/rotor seals are eliminated.
For temperatures above 2000 K, another problem becomes apparent. Heat transfer through radiation becomes significant. Highly turbulent gas flow also promotes heat transfer through radiation. As a consequence, even mono-crystal blades reach their melting temperature if air is used for cooling. Conventional fuels can provide temperatures in excess of 2300 K, while hydrogen would provide more than 2500 K. Temperatures at turbine inlet on current operating engines fall below the former level by more than 600 K.
Traditional methods of increasing temperature at the turbine inlet have been substantially exhausted and a new approach, based on non-traditional concepts is clearly needed to alleviate the aforementioned disadvantages.
It is therefore an object of the present invention to provide a liquid cooling system for a gas turbine where the liquid cooling system is incorporated in the turbine rotor, and where the operation of such a liquid cooling system is possible due to relative motion between the stator and the rotor of the turbine.
It is another object of the present invention to provide a liquid cooling system for a gas turbine which includes a tank containing a cooling liquid and a coolant pump coupled to the tank to extract the cooling liquid therefrom and to promote the cooling liquid through the rotor shaft, rotor disks and rotor blades by means of forced recirculation of the coolant liquid actuated as a result of relative motion of the stator with regard to the rotor.
It is a still further object of the present invention to provide a liquid cooling system in gas turbines where the heat exchanger is incorporated either within a compressor drum or, as part of the last stages of the compressor unit, or after the compressor, and where the liquid returning from the disks and blades of the turbine is cooled before returning to the liquid tank, either by circulation through the ducts within the body of the heat exchanger, or through the blades of the heat exchanger, or as a further option, passing through blades of the compressor section.
In accordance with the teachings of the present invention, a liquid cooling system is incorporated in the turbine rotor of a gas turbine and includes a tank containing a cooling liquid and a coolant pump coupled to the tank to extract the cooling liquid therefrom and to forcefully recirculate the cooling liquid through the rotor shaft, rotor disks, and rotor blades. The cooling system of the present invention further includes a heat exchanger and a system of channels (intake and output) extending within the rotor shaft, rotor disks, and blades. The inlet of the heat exchanger is coupled to the output channels to receive a heated cooling liquid therefrom, and the outlet of the heat exchanger is coupled to the inlet of the tank to deliver the cooled liquid thereto. The coolant pump of the liquid cooling system is actuated as a result of relative motion of the stator with regard to the rotor of the turbine.
There are several additional embodiments of the present invention with regard to the design of the heat exchanger, in addition to the positioning of the heat exchanger within the system. The heat exchanger may be positioned inside compressor drum. Also, the heat exchanger may be incorporated into the compressor stage or after said stage, where the heat exchanger would include cooling blades having ducts formed therethrough and coupled to the output channels, in order that the heated liquid delivered from the rotor blades and rotor disks to the cooling blades"" ducts transfers the heat associated therewith to the cooling blades of the heat exchanger. When the heat exchanger is incorporated into the compressor, the compressor""s blades are also used for cooling purposes.
Preferably, the cooling blades of the heat exchanger are formed of heat resistant steel or super alloy (for the embodiment of the heat exchanger positioned at the end of the compression section), while the heat exchanger positioned within the compressor drum may be formed as a casting made from either a heat resistant steel, super alloy, or titanium alloy.
The cooling pump is a pump for forced recirculation of the coolant liquid through the system, which includes a rotating pump body and an adjusting unit coupled to the rotating pump body for adjusting the speed of the cooling pump.
The tank is a rotating tank, adapted for containing a coolant liquid such as a metallic alloy, which would preferably be a composition of 25% of sodium and 75% of potassium.